U.S. patent application number 09/903641 was filed with the patent office on 2003-01-16 for method of combustor cycle airflow adjustment.
Invention is credited to Kojovic, Aleksandar, Stuttaford, Peter John.
Application Number | 20030010036 09/903641 |
Document ID | / |
Family ID | 25417851 |
Filed Date | 2003-01-16 |
United States Patent
Application |
20030010036 |
Kind Code |
A1 |
Stuttaford, Peter John ; et
al. |
January 16, 2003 |
Method of combustor cycle airflow adjustment
Abstract
A method of combustor cycle air flow adjustment for a gas
turbine engine according to the present invention solves the
problem of a higher flame temperature in the combustor, thereby
affecting the emission levels when a heat-recuperated air flow
cycle is used to increase the compressed air temperature. In low
emission combustors this impact is severe because emission levels
are significantly dependent on the primary combustion zone flame
temperature. The method of the present invention includes a step of
changing a geometry of an air flow passage and thereby changing
distribution of a total air mass flow between an air mass flow for
combustion and an air mass flow for cooling in order to ensure that
flame temperature in a primary combustion zone of a combustor are
maintained substantially the same whether the gas turbine engine is
manufactured to operate as a simple air flow cycle engine or as a
heat-recuperated air flow cycle engine. In an embodiment of the
present invention, the changing of the geometry of the air flow
passage by changing the number and size of perforations in an
impingement cooling skin so that with minimal changes the
impingement cooling skin serves duel purposes both as a cooling
device for cooling the combustor wall and as a valve means for
combustor cycle air flow adjustment, which makes the method simple
and economical.
Inventors: |
Stuttaford, Peter John;
(Toronto, CA) ; Kojovic, Aleksandar; (Oakville,
CA) |
Correspondence
Address: |
Gregory LaPointe
Bachman & LaPointe
900 Chapel Street
Suite 1201
New Haven
CT
06510-2802
US
|
Family ID: |
25417851 |
Appl. No.: |
09/903641 |
Filed: |
July 13, 2001 |
Current U.S.
Class: |
60/772 ;
60/39.23 |
Current CPC
Class: |
F02C 7/08 20130101; F05D
2260/201 20130101; F23R 2900/00016 20130101; F23R 3/005 20130101;
F23R 2900/03044 20130101; F23R 3/26 20130101 |
Class at
Publication: |
60/772 ;
60/39.23 |
International
Class: |
F23R 003/26 |
Claims
I/we claim:
1. A method of combustor cycle airflow adjustment for a gas turbine
engine design, comprising a step of changing a geometry of an air
flow passage and thereby changing distribution of a total air mass
flow between an air mass flow for combustion and an air mass flow
for cooling to ensure that a flame temperature in a primary
combustion zone of a combustor is maintained substantially the same
whether the gas turbine engine operates as a simple air flow cycle
engine or as a heat-recuperated air flow cycle engine.
2. The method as claimed in claim 1 wherein the geometry of the air
passage is changed to decrease the air mass flow for combustion
when the gas engine is used in a heat-recuperated air flow cycle,
with respect to the air mass flow for combustion when the engine is
used as a simple air flow cycle engine.
3. The method as claimed in claim 2 wherein the changing of the air
flow passage geometry is achieved by changing a geometry of a
cooling air passage.
4. The method as claimed in claim 2 wherein the changing of the air
flow passage geometry is achieved by changing a geometry of an
impingement skin of the combustor.
5. The method as claimed in claim 2 wherein the changing of the air
flow geometry is achieved by changing the number and size of a
plurality of small holes in an impingement skin of the combustor
such that the impingement skin performs as both a cooling device
and a valve for air mass flow distribution adjustment between the
simple flow cycle and the heat-recuperated flow cycle.
6. The method as claimed in claim 5 wherein the impingement skin
comprises a first group of holes therein adjacent to a combustor
wall section defining the primary combustion zone, the number and
size of the holes of the first group being predetermined to
substantially meet a cooling requirement for the combustor; and a
second group of holes therein adjacent to the a combustor wall
section defining a secondary combustion zone, the number and size
of the holes of the second group being adjusted to substantially
meet a flow distribution requirement when the air flow cycle of the
gas turbine engine is changed between the simple flow cycle and the
heat-recuperated flow cycle.
7. An impingement cooling skin in combination with a gas turbine
engine combustor to which the impingement cooling skin is attached,
the impingement skin comprising a first group of holes therein
adjacent to a combustor wall section defining a primary combustion
zone, the number and size of the holes of the first group being
predetermined to substantially meet a cooling requirement for the
combustor; and a second group of holes therein adjacent to a
combustor wall section defining a secondary combustion zone, the
number and size of the holes of the second group being adjusted to
substantially meet a flow distribution requirement when an air flow
cycle of the gas turbine engine is changed between a simple air
flow cycle and a heat-recuperated air flow cycle.
Description
FIELD OF THE INVENTION
[0001] The present invention relates to gas turbine engines,
particularly to air flow distribution adjustment of gas turbine
engines to ensure that low emissions not be affected by air flow
cycle patterns of the engines whether the engine is designed as a
simple air flow cycle or a heat-recuperated air flow cycle.
BACKGROUND OF THE INVENTION
[0002] Industrial gas turbine engines are subject to increasingly
stringent emission requirements. In order to provide a marketable
power generation product, an engine producing the lowest possible
emissions is crucial. Emissions of nitrogen oxides (NO.sub.x) and
carbon monoxide (CO) must be minimized over specified engine
operating ranges. To achieve this low level of emissions the
combustion system requires the complete burning of fuel and air at
low temperatures.
[0003] Combustors that achieve low NO.sub.x emissions without water
injection are known as dry-low emissions (DLE) and offer the
prospect of clean emissions combined with high engine efficiency.
This technology relies on a high air content in the fuel/air
mixture. While low emissions are critical, it is also important to
minimize the cost of manufacturing and maintaining the combustion
system, if a viable product is to be realized.
[0004] With regard to another aspect of the gas turbine engine,
engine efficiency is always of critical concern. It is known that
gas turbine efficiency can be substantially increased by
recuperating heat from the engine exhaust. In a gas turbine engine
using a heat-recuperating air flow cycle, compressor air is passed
through a recuperator or heat exchanger attached to the engine
exhaust end before entering the combustor. The resulting higher
combustor inlet air temperature thus requires less fuel burn to
achieve the same exit gas temperature compared with a gas turbine
engine using a simple air flow cycle in which the compressor air
enters the combustor directly and unaltered. The result of the
heat-recuperated air flow cycle is significantly improved gas
turbine cycle efficiency. The change in combustor inlet air
temperature and result in a combustor fuel/air ratio shifting to
achieve the same exhaust gas temperature. In low emission
combustors the impact of fuel/air ratios shifting is the most
severe because emission levels are intensely dependent on primary
combustion zone fuel/air ratios. Adjusting the combustor geometry
will compensate for the effect of fuel/air ratio change between
combustors of gas turbine engines using simple and heat-recuperated
air flow cycles. However, it can be cost prohibitive to adapt a gas
turbine engine model which operates as a simple air flow cycle
design, to be operable as a heat-recuperated air flow cycle design,
or vice versa,
[0005] Therefore, there is a need to develop a method of engine
design to cost effectively overcome the effect of the combustor
fuel/air ratio changing between simple and recuperated air flow
cycles when adapting a gas turbine engine model to be operable as
either a simple or a heat-recuperated air flow cycle design while
achieving low emission levels.
SUMMARY OF THE INVENTION
[0006] One object of the present invention is to provide a method
of low emission engine design for cost effectively overcoming the
effect of a combustor fuel/air ratio change between simple and
recuperated air flow cycles of the combustor engine.
[0007] Another object of the present invention is to provide a
method of combustor cycle air flow adjustment for a gas turbine
engine to ensure that the flame temperature is maintained
substantially the same whether the gas turbine engine utilizes a
simple air flow cycle or a heat-recuperated air flow cycle.
[0008] A further object of the present invention is to provide an
impingement cooling skin for a gas turbine engine combustor which
serves dual purposes as a cooling device to cool the combustor wall
and as a valve to adjust distribution of a total air mass flow
between an air mass flow for combustion and an air mass flow for
cooling, thereby changing the fuel/air ratio to ensure that the
combustor flame temperature is maintained substantially the same
whether the gas turbine engine operates as a simple air flow cycle
engine or a heat-recuperated air flow cycle engine.
[0009] In accordance with one aspect of the present invention, a
method of combustor cycle air flow adjustment for a gas turbine
engine design is provided. The method comprises a step of changing
a geometry of an air flow passage and thereby changing distribution
of a total air mass flow between an air mass flow for combustion
and an air mass flow for cooling, in order to ensure that a flame
temperature in a primary combustion zone of a combustor is
maintained substantially the same whether the gas turbine engine
operates as a simple air flow cycle engine or a heat-recuperated
air flow cycle engine.
[0010] The geometry of the air flow passage is preferably changed
to decrease the air mass flow for combustion when the gas turbine
engine uses a heat-recuperated air flow cycle, compared with the
air mass flow for combustion when the gas turbine engine uses a
simple air flow cycle. It is preferable that the changing of the
air flow passage geometry is achieved by changing a geometry of an
impingement skin of the combustor.
[0011] In accordance with another aspect of the present invention,
an impingement cooling skin in combination with a gas turbine
engine combustor is provided. The impingement cooling skin is
attached to the combustor and comprises a first group of holes
therein adjacent to a combustor wall section defining a primary
combustion zone, the number and size of the holes of the first
group being predetermined to substantially meet a cooling
requirement for the combustor. The impingement cooling skin further
includes a second group of holes therein adjacent to a combustor
wall section defining a secondary combustion zone, the number and
size of the holes of the secondary group being adjusted to
substantially meet a flow distribution requirement whether the gas
turbine engine operates as a simple air flow cycle engine or a
heat-recuperated air flow cycle engine.
[0012] The present invention advantageously provides a cost
effective solution to overcome the effect of combustor fuel/air
ratio change between a simple air flow cycle and a heat-recuperated
air flow cycle of the gas turbine engine operation. Without any
added parts, the impingement cooling skin of the combustor,
according to the present invention, serves dual purposes as a
cooling device to cool the combustor wall and as a valve to adjust
the distribution of the total air mass flow between the air mass
flow for combustion and the air mass flow for cooling such that the
required air mass flow for maintaining the combustor flame
temperature substantially the same, is achieved regardless of the
engine air flow cycle pattern. Thus, low emissions of the gas
turbine engine are ensured.
[0013] Other advantages and features of the present invention will
be better understood with reference to a preferred embodiment
described hereinafter.
BRIEF DESCRIPTION OF THE DRAWINGS
[0014] Having thus generally described the nature of the present
invention, reference will now be made to the accompanying drawings,
by way of example showing a preferred embodiment, in which:
[0015] FIG. 1 is a schematical illustration of a gas turbine engine
operating in a simple air flow cycle;
[0016] FIG. 2 is a schematical illustration of a gas turbine engine
operating in a heat-recuperated air flow cycle;
[0017] FIG. 3 is a graphical representation of the relationship
between NO.sub.x emission levels and the flame temperature;
[0018] FIG. 4 is a diagram of a air flow distribution of a gas
turbine engine incorporating the present invention; and
[0019] FIG. 5 is an elevational side view of a gas turbine
combustor with an impingement cooling skin attached thereto
according to one embodiment of the present invention, a section of
the combustor being cut away to show the cross-section of the
combustor.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
[0020] A simple air flow cycle of a gas turbine engine operation is
schematically illustrated in FIG. 1, in which the gas turbine
engine, generally indicated at numeral 10, includes a housing 12
having an air inlet 14 and a gas exhaust outlet 16. Within the
housing 12 a compressor 18 is driven by a turbine 20 through a
shaft 22. A combustor 24 is provided to generate combustion gases
in order to rotate the turbine 20 and further rotate a power
turbine 25 which in turn drives a generator 26 through a shaft 28.
The air flows in the simple air flow cycle of the gas turbine
engine operation are indicated by arrows 30. Air entering the inlet
14 is compressed by the compressor 18 and the compressed air then
directly enters the combustor 24 for combustion. The combustion
gases discharged from the combustor 24 are directed through the
turbine 20 and the power turbine 25, and then the exhausted gases
are discharged from the gas exhaust outlet 16.
[0021] A heat-recuperated air flow cycle of a gas turbine engine
operation is schematically illustrated in FIG. 2, in which the
engine, generally indicated at numeral 11 is similar to the gas
turbine engine 10 of FIG. 1. The gas turbine engine 11 includes
parts similar to those of gas turbine engine 10, which are
indicated by similar numerals and will therefore, not be
redundantly described herein. Additionally, gas turbine engine 11
includes a recuperator 32 which is a heat exchanger attached to the
gas exhaust outlet 16 to recover heat from the exhausted gases. Air
flows in the heat-recuperated air flow cycle of the gas turbine
engine operation are indicated by arrows 34. The air entering the
air inlet 14 is compressed by the compressor 18. However, instead
of directly flowing into the combustor 24, the compressed air is
directed to the recuperator 32 for heat exchange and then enters
the combustor 24. Similar to the simple air flow cycle shown in
FIG. 1, the combustion gases are discharged from the combustor 24
to rotate the turbine 20 and the power turbine 25 and then the
exhausted gases are discharged from the gas exhaust outlet 16. The
recuperator 32 transfers heat from the exhausted gases in the gas
exhaust outlet 16 to the compressed air and therefore, the
temperature of the compressed air entering the combustor 24 is much
higher than the temperature of the compressed air which enters the
combustor 24 in the simple air flow cycle as schematically
illustrated in FIG. 1. Higher compressed air temperatures make the
combustion operation more efficient, since less fuel is burned to
achieve the same exhaust gas temperature.
[0022] Nevertheless, lower compressed air temperatures increase the
amount of fuel to be burned and the fuel/air ratio in the combustor
24 becomes richer. It is already known that No.sub.x emissions
depend on flame temperatures as illustrated in FIG. 3, and it is
desirable to maintain the flame temperature at a low point to meet
the low NO.sub.x emission requirements. In a DLE system, the low
flame temperature point is achieved by maintaining a very lean
fuel/air ratio in the combustor, particularly in the primary
combustion zone. Generally the lean fuel/air ratio in the combustor
cannot be maintained when a gas turbine engine design is altered
from that of a simple air flow cycle operation to that of a
heat-recuperated air flow cycle operation, because more fuel must
be burned to achieve the same exhaust gas temperature. Adjusting
the combustor geometry will compensate for the effects of the
switch between simple and heat-recuperated air flow cycles of a gas
turbine engine operation. However, changing the combustor geometry
is cost prohibitive on engine designs when it is desired to adapt a
gas turbine engine model which operates as a simple air flow cycle
design, to be operable as a heat-recuperated air flow cycle design,
or vice versa.
[0023] A method of combustor cycle air flow adjustment for a gas
turbine engine is therefore developed according to the present
invention, and is generally illustrated in FIG. 4 in which a total
air mass flow 36 of a gas turbine engine is distributed between an
air mass flow 38 for a combustion system 42 of the engine and an
air mass flow 40 for a cooling system 44 of the same engine. The
exhausted combustion gases with exhausted cooling air are
discharged from the engine as indicated by arrow 46. The
temperature of the exhausted combustion gases must remain constant,
independent of cycle. 1f resistance to either air mass flow 38 or
40 is changed, the distribution of the total air mass flow 36
between the air mass flow 38 for the combustion system 42 and the
air mass flow 40 for the cooling system 44 will be changed
accordingly. The change of air mass flow resistance can be achieved
with a valve means 48 which may be connected to either air flow
branch, connected to the air flow branch for the air mass flow
40.
[0024] According to the present invention there is no need for the
valve means 48 physically presented herein, to adjust the air flow
resistance in either air flow branch. Changing the geometry of an
air passage for either air mass flow 38 or 40 will perform a
virtual valve function, regulating the air flow resistance of the
air passage for the air mass flow 38 or 40, thereby resulting in
adjustment of the air flow distribution. When the engine is to be
designed for a heat-recuperated air flow cycle operation and that
design is adapted from an engine design for a simple air flow cycle
operation, changing the geometry of the air passage either for the
air mass flow 38 or the air mass flow 40 in order to decrease the
air mass flow 38, will compensate for the decreased fuel burn,
thereby maintaining the flame temperature in the primary combustion
zone of the combustor and exhaust gas temperature, substantially
the same whichever original engine design is used.
[0025] In accordance with one embodiment of the present invention,
it is convenient and cost effective to change the geometry of an
impingement cooling skin to affect the air mass flow 38 for the
combustion system 42 which is illustrated in FIG. 5. A cyclone
combustor 110 includes a cylindrical combustor can 112 having a
central axis 114, an upstream end 116 and a downstream end 118
defined by an annular side wall 120. The upstream end 116 is closed
by an upstream end wall 122 and the downstream end 118 is in fluid
communication with a turbine section of the engine (not shown).
Three entry openings 124 (only one is shown) are provided in the
annular side wall 120 adjacent to the upstream end wall 122 for
receiving premixed fuel/air mixture into the combustor can 112. The
combustion processing of the premixed fuel/air mixture takes place
generally in a primary combustion zone 126 which is defined within
an upstream section of the combustor can 112. The combustion
products generated within the primary combustion zone 126, as well
as the un-reacted fuel and air will complete the combustion process
in a secondary combustion zone 128 which is a section of the
combustor can 112 downstream of the primary combustion zone 126.
The final combustion products are then discharged from the
downstream end 118 into the turbine.
[0026] Three fuel and air premixing tubes 130, such as venturi
premixing tubes,(only two are shown) are attached to the side wall
120 of the combustor can 112 and are positioned adjacent to the
upstream end wall 122. The premixing tubes 130 are
circumferentially, equally spaced apart from one another and are in
fluid communication with the combustor can 112 through the
respective entry openings 124 in the side wall 120.
[0027] Each premixing tube 130 includes a major tube section 132
for producing the fuel/air mixture therein and an outlet section
134 for injecting the fuel/air mixture into the combustor can 112
for combustion. The major tube section 132 has a central axis 136
thereof extending substantially parallel to the central axis 114 of
the combustor can 112. The outlet section 134 has a central axis
138 thereof extending substantially perpendicular to the central
axis 136 of the major tube section 132 and is oriented toward the
combustor can 112 radially with a tangential offset. Thus, lean
fuel/air mixture flows injected from the respective entry openings
124 in the side wall 120 create a swirling helical pattern within
the primary combustion zone 126 of the combustor can 112 as a
result of the tangential offset of the lean fuel/air mixture flows
exiting from the outlet sections 134 of the premixing tubes 130,
respectively. The swirling helical pattern of the burning lean
fuel/air mixture in the primary combustion zone 126 provides
optimum circulation of low temperature flames in the combustor can
112 which improves the liner life span of the combustor can 112,
flame stability in the combustion process and engine turn-down, as
well as the reduction of combustion noise and emission levels.
[0028] The size of the tube advantageously inhibits flashback from
the primary combustion zone 126 into the major tube section 132 of
the premixing tube 130. Pilot fuel lines 142 and ignitors 146 are
placed to take advantage of the locations of the entry openings 124
and the tangential direction of the lean fuel/air mixture flow
momentum generated from the tangential offset of the premixing
tubes 130.
[0029] An impingement cooling skin 148 is provided. The impingement
cooling skin 148 is made of a wrap-around sheet metal with
perforations therein and is positioned around the annular side wall
120 of the combustor can 112 in a radially spaced apart
relationship. It is optional that the impingement cooling skin 148
includes a perforated end skin 149 positioned axially spaced apart
from the upstream end wall 122 of the combustor can 112.
[0030] The perforations in the impingement cooling skin 148 as well
as in the end skin 149 are formed in two groups. The first group of
perforations includes holes 150 in the end skin 149 and in a
section of the impingement cooling skin 148 adjacent to a combustor
wall section defining the primary combustion zone 126. A second
group of perforations includes holes 152 in a section of the
impingement cooling skin 148 adjacent to a combustor wall section
defining the secondary combustion zone 128. The first group of
holes 150 are distributed in a predetermined pattern with a
predetermined number and a predetermined size thereof to
substantially meet the cooling requirements of the combustor wall
section defining the primary combustion zone 126. The size and
number, as well as the distribution pattern of the second group of
holes 152 are subject to change when the cyclone combustor 110 is
designed for different air flow cycle operations, in order to
substantially meet the flow distribution requirements relating to
the particular air flow cycle of a gas turbine engine operation, as
well as to provide secondary zone cooling.
[0031] The impingement cooling skin 148 further includes a sealing
ring 154 which is sealingly connected to a housing 156, only a
section of which is shown in FIG. 5. The housing 156 contains the
combustor can 112, and an annulus 158 is thereby formed
therebetween as a section of the air passage for cooling air. When
compressed air approaches the cyclone combustor 110 from above, the
total air mass flow is distributed between the air mass flow 38
which enters the premixing tubes 130 to mix with fuel to produce
the fuel/air mixture for combustion in the combustor can 112, and
the air mass flow 40 which enters holes 150 in the end skin 149 and
holes 150 and 152 in the impingement cooling skin 148 by way of the
annulus 158. The air entering the holes 150 in the end skin 149 and
holes 150 and 152 in the impingement cooling skin 148 impinges upon
the upstream end wall 122 and side wall 120 of the combustor can
112 to remove the heat therefrom. The combustor wall section
defining the primary combustion zone 126 is generally exposed to
higher temperatures than the combustor wall section defining the
secondary combustion zone 128, therefore the combustor wall section
defining the primary combustion zone 126 is cooled by the air from
the holes 150 which are specifically designed to substantially meet
the cooling requirements regardless of the air flow cycles of the
gas turbine engine operation.
[0032] In contrast to the predetermined pattern of the holes 150,
the holes 152 are designed substantially in accordance with the air
flow distribution requirements for the particular air flow cycles
of gas turbine engine operation. When the combustor 110 is designed
for a simple air flow cycle operation, the number and size of the
holes 152 are decreased to increase the air resistance of the
cooling air passage relative to the air mass flow distribution for
a heat-recuperated air flow cycle operation, such that less air
mass flow 40 is directed for cooling while relatively more air mass
flow 38 is directed to the premixing tubes 130 for combustion.
[0033] On the other hand, when the combustor 110 is designed for a
heat-recuperated air flow cycle the number and size of the air
holes 152 are increased to reducethe air resistance in the cooling
air passage such that relative to the air mass distribution for the
simple air flow cycle operation, more air mass flow 40 is directed
for cooling and relatively less air mass flow 38 is directed into
the premixing tubes 130 for combustion.
[0034] In the embodiment of the present invention, by using such a
method, the impingement cooling skin 148, 149 serves duel purposes
both as a cooling device to cool the combustor wall 120, 122 and as
a valve means to adjust air distribution flow between air mass
flows 38, 40 for combustion. Thus, it is convenient and cost
effective to just modify the impingement cooling skin 148, 149
design without need for other complicated changes of the combustor
in order to provide a gas turbine engine design for either simple
or heat-recuperated air flow cycle operation while maintaining
emissions at a same low level. Changing the number and size of
holes 152 of the impingement cooling skin 148 will also affect
cooling results. However, the holes 152 are designated for air
impingement upon the combustor wall section defining the secondary
combustion zone 128 which is exposed to lower temperatures and
therefore the negative effect of cooling is reduced. The cooling
aspect is optimized for the more severe conditions related to
heat-recuperated cycles.
[0035] Optionally, a gas turbine engine combustor may be
manufactured with a removable cover sheet or ring to selectively
cover the holes 152 of the impingement cooling skin 148, according
to a further embodiment of the present invention. Thus, the gas
turbine engine is adapted to operate alternatively as a simple or
heat-recuperated air flow cycle engine while maintaining low
emission levels.
[0036] The method of combustor cycle air flow adjustment for a gas
turbine engine according to the present invention is applicable to
various combustor systems, and the cyclone combustor described
above is an example only, illustrating a particular embodiment of
the present invention. This invention is also applicable to swirled
diffusion dump combustion designs. Modifications and improvements
to the above-described embodiment of the present invention may
become apparent to those skilled in the art. The foregoing
description is intended to be exemplary rather than limiting. The
scope of the invention is therefore intended to be limited solely
by the scope of the appended claims.
* * * * *